1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine blade with near wall cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as those used in electric power production with the industrial gas turbine engine, includes a turbine section with multiple rows or stages or turbine blades and vanes that react with a hot gas flow to drive the engine. The efficiency of the engine can be increased by passing a higher gas flow temperature into the turbine. However, the turbine inlet temperature is limited to the material properties of the first stage blades and vanes and the amount of cooling provided to them. Higher temperatures can be used with better cooling of the airfoils.
Besides higher engine efficiencies, designing the engine parts with longer life times also reduces the cost of operating the engine, a short lived part, such as a turbine blade, can significantly reduce the engine efficiency if the part is damaged enough to allow the hot gas flow to leak across the airfoil. Since an IGT engine operates continuously for years before a shut-down is scheduled, a damaged part might have to be used for long periods. It is very costly to shut down and engine and remove the parts for replacement. The down time means that the power plant operator does not have the use of the engine for power production. Also, warranties on the engine require the OEM (Original Equipment Manufacturer) to provide for a minimum period of time in which the engine must, run before a required shut-down is performed. Thus, part life is just as important as the part efficiency or performance in the engine.
In a typical turbine airfoil, such as a rotor blade or a stator vane, the airfoil is exposed to different temperatures. The leading edge is exposed to the highest temperature while the suction side near to the leading edge region is exposed to the lowest temperature. Thus, certain surfaces of the airfoil require more cooling than other surfaces. Also, the airfoil surface is exposed to different pressures. One surface of the airfoil would require a higher cooling air pressure than another area so that hot gas injection does not occur. If the hot gas flow can flow into the cooling air passages formed within the airfoil, significant damage can occur.
A prior art turbine blade with cooling passages is shown in FIG. 1. this prior art blade uses near wall cooling in the airfoil main body that is constructed of radial flow channels in the airfoil walls plus re-supply holes in conjunction with film discharge cooling holes. As a result of this cooling design, the spanwise and the chordwise cooling flow control due to the airfoil external hot gas temperature and pressure variation is difficult to achieve. Also, single radial channel flow is not the best method of utilizing cooling air and results in a low convective cooling effectiveness.